Engine assembly with exhaust pipe nozzle

ABSTRACT

An engine assembly including an internal combustion engine, an impulse turbine, and an exhaust pipe providing fluid communication between the exhaust port of the internal combustion engine and the flow path of the turbine. The exhaust pipe terminates in a nozzle. A ratio Vp/Vd between the pipe volume Vp and the displacement volume Vd of the internal combustion engine is at most 1.5. A minimum value of a cross-sectional area of the exhaust pipe is defined at the nozzle. In one embodiment, a ratio An/Ae between the minimum cross-sectional area An and the cross-sectional area Ae of the exhaust port of the internal combustion engine is at least 0.2. A method of compounding at least one internal combustion engine is also discussed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.14/740,889 filed Jun. 16, 2015, the entire contents of which areincorporated by reference herein.

TECHNICAL FIELD

The application relates generally to engine assemblies and, moreparticularly, to such engine assemblies including one or more internalcombustion engine(s).

BACKGROUND OF THE ART

Compound engine assemblies including an impulse turbine with areciprocating engine core are known and typically include long exhaustpipes with a relatively high pipe volume which is sufficient to maintainthe pressure of the exhaust pulses at an acceptable level correspondingto an acceptable flow speed at the connection with the impulse turbine.However, longer exhaust pipes typically lead to greater power lossesfrom the exhaust pulse and accordingly, less recovery in the compoundturbine.

SUMMARY

In one aspect, there is provided an engine assembly comprising: aninternal combustion engine having a working chamber of variable volume,the variable volume varying between a minimum volume and a maximumvolume with a difference between the maximum and minimum volumesdefining a displacement volume Vd; a turbine configured as an impulseturbine having a pressure based reaction ratio of at most 0.25, theturbine having a circumferential array of rotor blades extending acrossa flow path; and an exhaust pipe having a pipe volume Vp, the exhaustpipe providing fluid communication between an exhaust port of theinternal combustion engine and the flow path of the turbine, the exhaustpipe terminating in a nozzle communicating with a portion of the flowpath located upstream of the rotor blades; wherein a ratio Vp/Vd betweenthe pipe volume Vp and the displacement volume Vd of the internalcombustion engine is at most 1.5; and wherein a minimum value of across-sectional area of the exhaust pipe is defined at the nozzle.

In another aspect, there is provided an engine assembly comprising: aninternal combustion engine having a working chamber of variable volume,and having an inlet port and an exhaust port in communication with theworking chamber, the exhaust port having a cross-sectional area Ae; aturbine configured as an impulse turbine having a pressure basedreaction ratio of at most 0.25, the turbine having a circumferentialarray of rotor blades extending across a flow path; and an exhaust pipeproviding fluid communication between the exhaust port of the internalcombustion engine and the flow path of the turbine, the exhaust pipeterminating in a nozzle communicating with a portion of the flow pathlocated upstream of the rotor blades; wherein the nozzle includes aportion of reduced cross-sectional area with respect to a remainder ofthe exhaust pipe, the nozzle defining a minimum cross-sectional area Anof the exhaust pipe; and wherein a ratio An/Ae between the minimumcross-sectional area An and the cross-sectional area Ae of the exhaustport is at least 0.2.

In a further aspect, there is provided a method of compounding aninternal combustion engine, the method comprising: providing a turbineconfigured as an impulse turbine having a pressure based reaction ratioof at most 0.25; drivingly engaging the internal combustion engine andthe turbine to a common load; circulating exhaust gas from an exhaustport of the internal combustion engine through a first portion of anexhaust pipe having a first cross-sectional area and then through anozzle of the exhaust pipe having a second cross-sectional area smallerthan the first cross-sectional area; and circulating the exhaust gasfrom the nozzle to an inlet of the turbine, including directing theexhaust gas onto blades of the turbine.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a block diagram of a compound engine assembly according to aparticular embodiment;

FIG. 2 is a cross-sectional view of a Wankel engine which can be used ina compound engine assembly such as shown in FIG. 1, according to aparticular embodiment;

FIG. 3 is a schematic representation of part of the compound engineassembly of FIG. 1 according to a particular embodiment;

FIG. 4 is a schematic graph of pressure variation with respect to crankangle, for a compound engine including exhaust pipes with a constantcross-sectional area and for a compound engine including exhaust pipeswith nozzles as per a particular embodiment; and

FIG. 5 is a schematic graph of flow speed variation with respect tocrank angle, for the compound engine including exhaust pipes with aconstant cross-sectional area and for the compound engine includingexhaust pipes with nozzles.

DETAILED DESCRIPTION

Referring to FIG. 1, a compound engine assembly 10 is schematicallyshown. The compound engine assembly 10 includes an engine core with oneor more internal combustion engine(s) 12. The core engine(s) 12 drive acommon load. In the embodiment shown, the common load includes an outputshaft 16 which may be for example connected to a propeller through areduction gearbox (not shown) and to which each core engine 12 isengaged. Other possible common loads may include, but are not limitedto, one or more compressor and/or fan rotor(s), electrical generator(s),accessories, rotor mast(s), or any other type of load or combinationthereof.

In a particular embodiment, the compound engine assembly 10 alsoincludes a turbocharger 18, including a compressor 20 and a second stageturbine 22 which are drivingly interconnected by a shaft 24. In aparticular embodiment, the second stage turbine 22 is a pressureturbine, also known as a reaction turbine. The compressor 20 and thesecond stage turbine 22 may each be a single-stage device or amultiple-stage device with a single shaft or split on multipleindependent shafts in parallel or in series, and may each be acentrifugal or axial device. The compressor 20 of the turbocharger 18compresses the air before it enters the core engine(s) 12. Thecompressor 20 and the second stage turbine 22 may each include one ormore rotors, with radial, axial or mixed flow blades.

In the embodiment shown, the shaft 24 of the turbocharger 18 rotatesindependently of the common load. The turbocharger shaft 24 may extendsalong a different axis than that of the output shaft 16, for exampletransverse to the output shaft 16, or may be defined coaxially with theoutput shaft 16; the turbocharger shaft and output shaft 16 may belinked to rotate together, for example through a transmission, or mayrotate independently from one another.

Alternately, the turbocharger 18 may be omitted.

Each core engine 12 provides an exhaust flow in the form of exhaustpulses. The exhaust flow from the core engines 12 is supplied to acompound or first stage turbine 26 in fluid communication therewith,also driving the common load. The first stage turbine 26 is configuredas a velocity type turbine, also known as an impulse turbine, and couldbe an axial, radial or mixed flow turbine.

A pure impulse turbine works by changing the direction of the flowwithout accelerating the flow inside the rotor; the fluid is deflectedwithout a significant pressure drop across the rotor blades. The bladesof the pure impulse turbine are designed such that in a transverse planeperpendicular to the direction of flow, the area defined between theblades is the same at the leading edges of the blades and at thetrailing edges of the blade: the flow area of the turbine is constant,and the blades are usually symmetrical about the plane of the rotatingdisc. The work of the pure impulse turbine is due only to the change ofdirection in the flow through the turbine blades. Typical pure impulseturbines include steam and hydraulic turbines.

In contrast, a reaction turbine accelerates the flow inside the rotorbut needs a static pressure drop across the rotor to enable this flowacceleration. The blades of the reaction turbine are designed such thatin a transverse plane perpendicular to the direction of flow, the areadefined between the blades is larger at the leading edges of the bladesthan at the trailing edges of the blade: the flow area of the turbinereduces along the direction of flow, and the blades are usually notsymmetrical about the plane of the rotating disc. The work of the purereaction turbine is due mostly to the acceleration of the flow throughthe turbine blades.

Most aeronautical turbines are not “pure impulse” or “pure reaction”,but rather operate following a mix of these two opposite butcomplementary principles—i.e. there is a pressure drop across theblades, there is some reduction of flow area of the turbine blades alongthe direction of flow, and the speed of rotation of the turbine is dueto both the acceleration and the change of direction of the flow. Thedegree of reaction of a turbine can be determined using thetemperature-based reaction ratio (equation 1) or the pressure-basedreaction ratio (equation 2), which are typically close to one another invalue for a same turbine:

$\begin{matrix}{{{Reaction}(T)} = \frac{\left( {t_{S\; 3} - t_{S\; 5}} \right)}{\left( {t_{S\; 0} - t_{S\; 5}} \right)}} & (1) \\{{{Reaction}(P)} = \frac{\left( {P_{S\; 3} - P_{S\; 5}} \right)}{\left( {P_{S\; 0} - P_{S\; 5}} \right)}} & (2)\end{matrix}$

where T is temperature and P is pressure, s refers to a static port, andthe numbers refers to the location the temperature or pressure ismeasured: 0 for the inlet of the turbine vane (stator), 3 for the inletof the turbine blade (rotor) and 5 for the exit of the turbine blade(rotor); and where a pure impulse turbine would have a ratio of 0 (0%)and a pure reaction turbine would have a ratio of 1 (100%).

Aeronautical turbines referred to as impulse turbines typically have areaction ration of 0.25 (25% reaction) or lower, although other valuesare also possible.

In a particular embodiment, the first stage turbine 26 is configured totake benefit of the kinetic energy of the pulsating flow exiting thecore engine(s) 12 while stabilizing the flow, and the second stageturbine 22 is configured to extract energy from the remaining pressurein the flow. Accordingly, the first stage turbine 26 has a lower (i.e.lower value) reaction ratio than that of the second stage turbine 22.

In a particular embodiment, the second stage turbine 22 has a reactionratio higher than 0.25; in another particular embodiment, the secondstage turbine 22 has a reaction ratio higher than 0.3; in anotherparticular embodiment, the second stage turbine 22 has a reaction ratioof about 0.5; in another particular embodiment, the second stage turbine22 has a reaction ratio higher than 0.5.

In a particular embodiment, the first stage turbine 26 has a reactionratio of at most 0.2; in another particular embodiment, the first stageturbine 26 has a reaction ratio of at most 0.15; in another particularembodiment, the first stage turbine 26 has a reaction ratio of at most0.1; in another particular embodiment, the first stage turbine 26 has areaction ratio of at most 0.05.

It is understood that any of the above-mentioned reaction ratios for thesecond stage turbine 22 can be combined with any of the above-mentionedreaction ratios for the first stage turbine 26 and that these ratios canbe pressure-based or temperature-based. Other values are also possible.

In the embodiment shown, the first stage turbine 26 is connected to theoutput shaft 16 through an appropriate type of transmission 28, forexample a planetary, star, offset or angular gear system. The outlet ofthe first stage turbine 26 is in fluid communication with an inlet ofthe second stage turbine 22. Energy is extracted from the exhaust gasexiting the first stage turbine 26 by the second stage turbine 22 todrive the compressor 20 via the connecting shaft 24.

Although not shown, the air may optionally circulate through anintercooler between the compressor 20 and the core engine(s) 12, and thefirst stage engine assembly 10 also includes a cooling system, includingfor example a circulation system for a coolant (e.g. water-ethylene,oil, air) to cool the housing of each core engine 12, an oil coolant forthe internal mechanical parts of the core engine(s) 12, one or morecoolant heat exchangers, etc.

The fuel injector(s) of each core engine 12, which in a particularembodiment are common rail fuel injectors, communicate with a source 30of Heavy fuel (e.g. diesel, kerosene (jet fuel), equivalent biofuel),and deliver the heavy fuel into the core engine(s) 12 such that thecombustion chamber is stratified with a rich fuel-air mixture near theignition source and a leaner mixture elsewhere.

In a particular embodiment, each core engine 12 is a rotary internalcombustion engine having a rotor sealingly engaged in a respectivehousing. In a particular embodiment, the rotary engine(s) is/are Wankelengine(s). Referring to FIG. 2, an exemplary embodiment of a Wankelengine is shown; it is understood that the configuration of the coreengine(s) 12 used in the compound engine assembly 10, e.g. placement ofports, number and placement of seals, etc., may vary from that of theembodiment shown. In addition, it is understood that each core engine 12may be any other type of internal combustion engine including, but notlimited to, any other type of rotary engine.

As shown in FIG. 2, each Wankel engine comprises a housing 32 definingan internal cavity with a profile defining two lobes, which ispreferably an epitrochoid. A rotor 34 is received within the internalcavity. The rotor defines three circumferentially-spaced apex portions36, and a generally triangular profile with outwardly arched sides. Theapex portions 36 are in sealing engagement with the inner surface of aperipheral wall 38 of the housing 32 to form three working chambers 40between the rotor 34 and the housing 32.

The rotor 34 is engaged to an eccentric portion 42 of the output shaft16 to perform orbital revolutions within the internal cavity. The outputshaft 16 performs three rotations for each orbital revolution of therotor 34. The geometrical axis 44 of the rotor 34 is offset from andparallel to the axis 46 of the housing 32. During each orbitalrevolution, each chamber 40 varies in volume and moves around theinternal cavity to undergo the four phases of intake, compression,expansion and exhaust. The difference between the maximum and minimumvolumes of each chamber 40 during the revolutions of the rotor 34defines a displacement volume Vd of the engine.

An intake port 48 is provided through the peripheral wall 38 forsuccessively admitting compressed air into each working chamber 40. Anexhaust port 50 is also provided through the peripheral wall 38 forsuccessively discharging the exhaust gases from each working chamber 40.Passages 52 for a glow plug, spark plug or other ignition element, aswell as for one or more fuel injectors (not shown) are also providedthrough the peripheral wall 38. Alternately, the intake port 48, theexhaust port 50 and/or the passages 52 may be provided through an end orside wall 54 of the housing; and/or, the ignition element and a pilotfuel injector may communicate with a pilot subchamber (not shown)defined in the housing 32 and communicating with the internal cavity forproviding a pilot injection. The pilot subchamber may be for exampledefined in an insert (not shown) received in the peripheral wall 38.

For efficient operation the working chambers 40 are sealed, for exampleby spring-loaded apex seals 56 extending from the rotor 34 to engage theperipheral wall 38, and spring-loaded face or gas seals 58 and end orcorner seals 60 extending from the rotor 34 to engage the end walls 54.The rotor 34 also includes at least one spring-loaded oil seal ring 62biased against the end wall 54 around the bearing for the rotor 34 onthe shaft eccentric portion 42.

Each Wankel engine provides an exhaust flow in the form of a relativelylong exhaust pulse; for example, in a particular embodiment, each Wankelengine has one explosion per 360° of rotation of the output shaft, withthe exhaust port remaining open for about 270° of that rotation, thusproviding for a pulse duty cycle of about 75%. By contrast, a piston ofa reciprocating 4-stroke piston engine typically has one explosion per720° of rotation of the output shaft with the exhaust port remainingopen for about 180° of that rotation, thus providing a pulse duty cycleof 25%.

In a particular embodiment which may be particularly but not exclusivelysuitable for low altitude, each Wankel engine has a volumetric expansionratio of from 5 to 9, and a volumetric compression ratio lower than thevolumetric expansion ratio. The power recovery of the first stageturbine 26 may be maximized by having the exhaust gas temperatures atthe material limit, and as such is suitable for such relatively lowvolumetric compression ratios, which may help increase the power densityof the Wankel engine and may also improve combustion at high speed andof heavy fuel.

Referring to FIG. 3, in a particular embodiment, the compound engineassembly 10 includes two (2) core engines 12 in the form of Wankelengines, for example such as shown in FIG. 2, with the two eccentricportions 42 of the output shaft 16 being angularly offset at 180° fromone another for balancing of the compound engine assembly 10. In otherembodiments, more or less core engines 12 may be provided; for example,in another particular embodiment, the core includes four (4) Wankelengines.

In the embodiment shown, the transmission 28 of the first stage turbine26 includes a sun gear 76 attached on the shaft of the rotor of thefirst stage turbine 26, and an array of planet gears 78 meshed with thesun gear 76. The planet gears 78 are mounted on a rotating carrier whichis drivingly engaged to the output shaft 16. The planet gears 78 aremeshed with a stationary ring gear 79. In another embodiment, the planetgears 78 are mounted on a stationary carrier, and are meshed with a ringgear drivingly engaged to the output shaft 16. The speed reduction ratioof the transmission 28 may be selected to optimize operation of thefirst stage turbine 26 and of the core engines 12. Other configurationsare also possible. For example, the first stage turbine 26 may bemounted in an offset manner rather than co-axially with the core engines12. The first stage turbine 26 may be drivingly engaged to the outputshaft 16 through an angular, for example perpendicular, transmissionsystem, for example including a gearbox and a tower shaft.

The rotor blades 64 of the first stage turbine 26 extend across anannular flow path 66. In the embodiment shown, the rotor of the firststage turbine 26 is an axial rotor and the flow path 66 extends axially.Although not shown, in all embodiments, variable geometry elements suchas inlet guide vanes, blow-off valves, waste gates, etc. may be used toobtain desired system operability.

The assembly 10 includes an exhaust pipe 68 for each core engine 12.Each exhaust pipe 68 extends from the respective exhaust port 50 (seealso FIG. 2) of the respective core engine 12 to a portion of the firststage turbine flow path 66 located upstream of the rotor blades 64, toprovide the fluid communication therebetween.

The flow path 66 and/or the outlet of each exhaust pipe 68 are shaped todirect the exhaust pulses onto the blades 64 to allow the exhaust pulsesto drive rotation of the rotor of the first stage turbine 26. Theexhaust pipes 68 extend independently from one another, and have arelatively small length, which in a particular embodiment allows forincreased use of the exhaust pulse kinetic energy to drive the firststage turbine 26. In a particular embodiment, the length of the exhaustpipes 68 is small enough such that the ratio Vp/Vd between the internalvolume Vp of each exhaust pipe 68 and the displacement volume Vd of thecorresponding core engine 12 is at most 1.5; in another particularembodiment, the ratio Vp/Vd is at most 1.2; in another particularembodiment, the ratio Vp/Vd is at most 1.0.

Shorter exhaust pipes 68 can lead to high pressures at the exhaust pipeoutlet; a higher flow pressure cause a higher density of the flow, whichleads to a lower flow speed; lower flow speed may in turn lead to lowerenergy recovery in the first stage turbine 26. In a particularembodiment, this effect is compensated by including a nozzle 70 in theportion of each exhaust pipe 68 defining the communication with thefirst stage turbine flow path 66. In the embodiment shown, the nozzle 70is located at the outlet end of the exhaust pipe 68, immediatelyupstream of the flow path 66. The nozzle 70 defines the minimalcross-sectional area of each exhaust pipe 68. The reduced cross-sectionof the exhaust pipe 68 defined by the nozzle 70 allows for the flowspeed through the nozzle 70 to be increased, which in a particularembodiment allows to obtain improved energy recovery in the first stageturbine 26 with respect to an exhaust pipe 68 of similar length andvolume but with a constant cross-section throughout its entire length.

In the embodiment shown, the nozzles 70 and accordingly thecommunications between the exhaust pipes 68 and the first stage turbineflow path 66 are spaced apart around the circumferential direction ofthe first stage turbine 26.

In the embodiment shown, the cross-sectional area of each exhaust pipe68 is constant upstream of the nozzle 70, for example in the portion 69of the exhaust pipe 68 extending from its communication with the exhaustport 50 to the nozzle 70, and this constant cross-sectional areacorresponds to that of the engine exhaust port 50 connected to theexhaust pipe 68. Alternately, the exhaust pipe 68 may have a differentcross-sectional area than that of the exhaust port 50, and may be forexample smaller than that of the corresponding exhaust port 50.

In a particular embodiment, the ratio An/Ae between the minimum value Anof the cross-sectional area of each exhaust pipe 68 at the nozzle 70 andthe cross-sectional area Ae of the corresponding exhaust port 50 is atleast 0.2. In other embodiments, the ratio An/Ae may be at least 0.4, atmost 0.6, from 0.2 to 0.6, or from 0.4 to 0.6. In a particularembodiment, the ratio An/Ae is about 0.5. Other values are alsopossible.

In addition, an inlet pipe 72 is connected to each intake port 48 (seealso FIG. 2), and if a turbocharger is provided, provides fluidcommunication between the intake port 48 and the outlet of thecompressor 20 (FIG. 1). A turbine pipe 80 extends from the flow path 66of the first stage turbine 26 downstream of the rotor blades 64, and ifa turbocharger is provided, provides fluid communication between thefirst stage turbine 26 exhaust and the inlet of the second stage turbine22 (FIG. 1).

Accordingly, in a particular embodiment, the internal combustionengine(s) 12 of the engine core are compounded by providing the firststage turbine 26, drivingly engaging each engine 12 and the first stageturbine 26 to a common load, circulating the exhaust gas from eachexhaust port 50 through a first portion 69 of the respective exhaustpipe 68, for example having the same cross-sectional area Ae as that ofthe exhaust port 50, and then through the nozzle 70 having across-sectional area An smaller than that of the first portion 69. Theexhaust gas is then circulated from the nozzle to the inlet of the firststage turbine 26, and directed onto the blades 64 of its rotor.

FIGS. 4-5 show examples of the effect of the presence of the nozzle 70(reduced cross-sectional area) at the outlet of the exhaust pipe 68. Thetwo graphs show the comparison between two assemblies similar to theassembly 10 described above, the assemblies being identical aside fromthe configuration of their exhaust pipes 68. Both assemblies includeexhaust pipes defining a ratio Vp/Vd of 0.8. In the first assembly, theexhaust pipes include a nozzle at their outlet end defining a ratioAn/Ae of 0.95, i.e. the cross-sectional area of the exhaust pipes is notsubstantially reduced at the connection with the first stage turbineflow path 66. In the second assembly, the exhaust pipes include a nozzleat their outlet end defining a ratio An/Ae of 0.48.

In FIG. 4, the pressure variation with respect to the crank angle of theoutput shaft 16 is shown. It can be seen that the static pressure pulseP₁ of the assembly with the larger exit nozzles (An/Ae of 0.95) is lowerthan the static pressure pulse P₂ of the assembly with the smaller exitnozzles (An/Ae of 0.48); with the larger nozzles, more flow exits thepipe at the beginning of the pulse because of the larger cross-sectionalarea, and accordingly the pressure build up is less steep.

In FIG. 5, the variation in flow speed (Mach number—MN) at the outlet ofthe exhaust pipe adjacent the turbine inlet or communication with thefirst stage turbine flow path 66 with respect to the crank angle of theoutput shaft 16 is shown. It can be seen that the time during which thespeed of the flow is at the maximum value M_(MAX) (which in a particularembodiment corresponds to Mach 1) is longer for the assembly with thesmaller exit nozzles (An/Ae of 0.48), as shown by M₂, than for theassembly with the larger exit nozzles (An/Ae of 0.95), as shown by M₁.Thus, the presence of the reduced cross-sectional area as defined by thesmaller nozzles 70 allows for an increase in mean flow speed at thefirst stage turbine inlet. In a particular embodiment, the flow speed atthe turbine inlet M₂ of the assembly including the smaller exit nozzles(An/Ae of 0.48) reaches a maximal value M_(MAX) of about Mach 1, and amean value of about Mach 0.8. Other values are also possible. In aparticular embodiment, the larger mean value of flow speed provides fora better performance of the first stage turbine 26, and the largerportion of the full 360 degrees crank angle cycle at the maximal valueMn=1 allows for the transient effects of the pulsating flow in theturbine to be less disturbing and closer to a constant flow turbine,which may yield better efficiency.

In a particular embodiment, the nozzles 70 create equivalent or higherflow speeds at the outlet ends of short exhaust pipes than that obtainedat the outlet ends of long exhaust pipes which are sufficiently long toavoid pressure increases leading to flow speed reductions, whileavoiding the energy losses associated with the dampening of the exhaustpulses occurring across the length of such long exhaust pipes.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Modifications which fall within the scope of the present invention willbe apparent to those skilled in the art, in light of a review of thisdisclosure, and such modifications are intended to fall within theappended claims.

1. An engine assembly comprising: an internal combustion engine having aworking chamber of variable volume, the variable volume varying betweena minimum volume and a maximum volume with a difference between themaximum and minimum volumes defining a displacement volume Vd; a turbineconfigured as an impulse turbine having a pressure based reaction ratioof at most 0.25, the turbine having a circumferential array of rotorblades extending across a flow path; and an exhaust pipe having a pipevolume Vp, the exhaust pipe providing fluid communication between anexhaust port of the internal combustion engine and the flow path of theturbine, the exhaust pipe terminating in a nozzle communicating with aportion of the flow path located upstream of the rotor blades; wherein aratio Vp/Vd between the pipe volume Vp and the displacement volume Vd ofthe internal combustion engine is at most 1.5; and wherein a minimumvalue of a cross-sectional area of the exhaust pipe is defined at thenozzle.
 2. The engine assembly as defined in claim 1, wherein theturbine and the internal combustion engine are in driving engagementwith a common load.
 3. The engine assembly as defined in claim 1,wherein the ratio Vp/Vd is at most 1.0.
 4. The engine assembly asdefined in claim 1, wherein the cross-sectional area of the exhaust pipeupstream of the nozzle is constant.
 5. The engine assembly as defined inclaim 4, wherein the cross-sectional area of the exhaust pipe upstreamof the nozzle corresponds to a cross-sectional area of the exhaust portof the internal combustion engine.
 6. The engine assembly as defined inclaim 1, wherein a ratio An/Ae between the minimum value An of thecross-sectional area at the nozzle and a cross-sectional area Ae of theexhaust port of the internal combustion engine is at least 0.2.
 7. Theengine assembly as defined in claim 5, wherein the ratio An/Ae is atmost 0.6.
 8. The engine assembly as defined in claim 5, wherein theratio An/Ae is at least 0.4.
 9. The engine assembly as defined in claim1, wherein the turbine is a first stage turbine, the assembly furthercomprising a turbocharger including a compressor and a second stageturbine in driving engagement with one another, an outlet of thecompressor being in fluid communication with an inlet port of theinternal combustion engine, and an inlet of the second stage turbinebeing in fluid communication with a portion of the flow path of thefirst stage turbine located downstream of the rotor blades of the firststage turbine, the second stage turbine having a pressure based reactionratio higher than the reaction ratio of the first stage turbine.
 10. Theengine assembly as defined in claim 1, wherein further comprising anadditional internal combustion engine and an additional exhaust pipeproviding fluid communication between an exhaust port of the additionalinternal combustion engine and the flow path of the turbine, the nozzleof the exhaust pipe and the nozzle of the additional exhaust pipe beingspaced apart from each other along a circumferential direction of theturbine.
 11. The engine assembly as defined in claim 1, furthercomprising a heavy fuel source in communication with the internalcombustion engine.
 12. An engine assembly comprising: an internalcombustion engine having a working chamber of variable volume, andhaving an inlet port and an exhaust port in communication with theworking chamber, the exhaust port having a cross-sectional area Ae; aturbine configured as an impulse turbine having a pressure basedreaction ratio of at most 0.25, the turbine having a circumferentialarray of rotor blades extending across a flow path; and an exhaust pipeproviding fluid communication between the exhaust port of the internalcombustion engine and the flow path of the turbine, the exhaust pipeterminating in a nozzle communicating with a portion of the flow pathlocated upstream of the rotor blades; wherein the nozzle includes aportion of reduced cross-sectional area with respect to a remainder ofthe exhaust pipe, the nozzle defining a minimum cross-sectional area Anof the exhaust pipe; and wherein a ratio An/Ae between the minimumcross-sectional area An and the cross-sectional area Ae of the exhaustport is at least 0.2.
 13. The engine assembly as defined in claim 12,wherein the turbine and the internal combustion engine are in drivingengagement with a common load.
 14. The engine assembly as defined inclaim 12, wherein the cross-sectional area of the exhaust pipe upstreamof the nozzle is constant.
 15. The engine assembly as defined in claim14, wherein the cross-sectional area of the exhaust pipe upstream of thenozzle corresponds to the cross-sectional area Ae of the exhaust port.16. The engine assembly as defined in claim 12, wherein the ratio An/Aeis at most 0.6.
 17. The engine assembly as defined in claim 16, whereinthe ratio An/Ae is at least 0.4.
 18. The engine assembly as defined inclaim 12, wherein the turbine is a first stage turbine, the assemblyfurther comprising a turbocharger including a compressor and a secondstage turbine in driving engagement with one another, an outlet of thecompressor being in fluid communication with the inlet port of theinternal combustion engine, and an inlet of the second stage turbinebeing in fluid communication with a portion of the flow path of thefirst stage turbine located downstream of the rotor blades of the firststage turbine, the second stage turbine having a pressure based reactionratio higher than the reaction ratio of the first stage turbine.
 19. Theengine assembly as defined in claim 12, further comprising a heavy fuelsource in communication with the internal combustion engine.
 20. Amethod of compounding an internal combustion engine, the methodcomprising: providing a turbine configured as an impulse turbine havinga pressure based reaction ratio of at most 0.25; drivingly engaging theinternal combustion engine and the turbine to a common load; circulatingexhaust gas from an exhaust port of the internal combustion enginethrough a first portion of an exhaust pipe having a firstcross-sectional area and then through a nozzle of the exhaust pipehaving a second cross-sectional area smaller than the firstcross-sectional area; and circulating the exhaust gas from the nozzle toan inlet of the turbine, including directing the exhaust gas onto bladesof the turbine.
 21. The method as defined in claim 20, wherein a ratioVp/Vd between a pipe volume Vp of the exhaust pipe and a displacementvolume Vd of the internal combustion engine is at most 1.5.
 22. Themethod as defined in claim 20, wherein a minimum value of across-sectional area of the exhaust pipe is defined at the nozzle. 23.The method as defined in claim 20, wherein the internal combustionengine is a rotary engine.